Composite sandwich structures constructed of honeycomb core or other lightweight core materials provide several advantages over other composite structural arrangements. Sandwich structures typically include the core material bounded on opposing sides of the core by face sheets or laminates comprised of one or more plies of composite material. Due to the relatively light weight of the core, the combination of the core and laminates on opposing sides of the core results in a relatively high stiffness-to-weight ratio as compared to composite structures comprised of laminated plies. In addition, composite sandwich structures have relatively high strength-to-weight ratios due to the relatively low density of the core material. Certain structures such as, without limitation, wing flaps and/or doors of a commercial airliner may benefit from a composite sandwich construction due to the favorable stiffness characteristics and light weight.
During the fabrication of a composite structure, pressure and heat are typically applied to a layup of composite materials that make up the structure in order to cure and bond the composite materials. An autoclave may be employed as a means for applying heat and pressure to the composite material layup such as a composite sandwich structure layup. For the above-noted example of a wing flap formed with lightweight core material, the geometry and size of the flap may present challenges regarding the curing and bonding of the composite materials to the core.
For example, the wing flap of a commercial airliner may have relatively large core thicknesses (e.g., six inches or larger) with a trend toward increasing core thicknesses in response to ongoing efforts to minimize weight in composite structures. The wing flap may also include one or more chamfers for tapering the core thickness such as along a direction toward the perimeter of the flap. The chamfer angle may exceed 10 degrees and may be as large as 20 degrees or greater depending upon the flap geometry. For such a configuration, the pressure exerted by the autoclave (e.g., approximately 45 psi) may result in the application of a relatively large side load (e.g., up to 4,000 pounds) on the chamfer of the core. The relatively large side load may result in movement or slippage of the plies of the upper and lower laminates relative to one another along a direction from the flap perimeter toward the core chamfer. Furthermore, the heating of the plies between which the core is sandwiched may reduce the viscosity of the resin that is in contact with the core and which may reduce friction and further facilitate ply slippage. The slippage of the plies relative to one another may result in movement of the core causing the core to be compacted or crushed in response to the ply movement.
Prior art attempts to prevent core crush include the application of film adhesive and a fiberglass sheet around the border of the composite panel in an effort to stabilize the core from movement during the application of autoclave pressure. Unfortunately, the fiberglass sheet extends across the surface of the core resulting in additional weight to the structure. Other attempts to prevent core crush include the application of tie down straps along the perimeter of the composite layup. The tie down straps may comprise fiberglass straps that may be secured to the tool and extended over and adhered to the uppermost ply of the composite layup. However, such tie down straps may be ineffective against the relatively large side forces exerted on chamfers of large surface area during the application of autoclave pressure.
Another approach to preventing ply slippage in composite structures is to septumize or split the core along a horizontal plane at an approximate mid-height of the core. Layers of fiberglass and adhesives may be installed between upper and lower portions of the split core in an attempt to stabilize the core against movement. Unfortunately, the addition of the fiberglass and adhesive layers may eliminate the ability to perform a non-destructive inspection of the composite panel using ultrasonic inspection techniques due to blockage of the ultrasonic signal by the fiberglass layer. Furthermore, the addition of the fiberglass sheet and adhesive may add to the weight of the composite structure.
As can be seen, there exists a need in the art for a system and method for stabilizing the plies of a composite structure against movement relative to one another in order to prevent core crush during the application of pressure to the composite structure as may occur during curing and/or consolidation of the structure. Furthermore, there exists a need in the art for a system and method for stabilizing the plies of a composite structure that is effective for relatively thick cores having a chamfer formed at a relatively steep chamfer angle. Finally, there exists a need in the art for a system and method for stabilizing the plies of a composite structure against movement without requiring the addition of materials that may increase the weight of the composite structure.